Course-changing gun-launched missile



March 26, 1968 H. J. PLUMLEY 3,374,967

COURSECHANGING GUN-LAUNCHED MISSILE Filed DEC. 6. 1949 I 3 Sheets-Sheetl FIRE CONTROL March 26, 1968 H. J. PLUMLEY CHANGING GUN-LAUNCHEDMISSILE COURSE- Filed Dec. 6. 1949 5 Sheets-Sheet 7;

O a t March 26, 1968 H. J. PLUMLEY COURSECHANGING GUN-LAUNCHED MISSILE 3Sheets-Sheet 3 AMPL IF If R Filed Dec.

7'4 R6 E T INTERCEPT/ON RANGE 360 DEGREES 0F RQTAT/OA/ United StatesPatent I 3,374,967 COURSE-CHANGING GUN-LAUNCHED MISSILE Harold J.Plumley, Washington, D.C., assignor to the United States of America asrepresented by the Secretary of the Navy Filed Dec. 6, 1949, Ser. No.131,441 5 Claims. (Cl. 2443.14)

The invention described herein may be manufactured and used by or forthe Government of the United States of America for governmental purposeswithout the payment of any royalties thereon or therefor.

The present invention relates to apparatus for changing the course of agun-fired missile during the flight thereof. More particularly, theinvention relates to apparatus for changing the course of a rotating,fin-stabilized missile during the flight thereof by firing a reactionsteering charge contained in the missile through one side thereof and ina direction at right angles to the line of flight, the steering chargebeing arranged to direct its force through the center of gravity of themissile.

As a result of the steering force being directed through the center ofgravity of the missile, the missile is thrust laterally of the initialtrajectory and assumes the course of the new trajecory with itslongitudinal axis disposed at an angle thereto, the longitudinal axis ofthe missile being parallel to the initial trajectory. The stabilizingfins of the missile, however, serve to bring the longitudinal axis intoalignment with the new trajectory.

In aiming a missile toward a moving target it is a well known and easilysolved problem to compute the proper lead and angle of trajectory whenthe target is following a known straight line course at a known speed.Where the target changes its course and/or speed after an uncontrolledmissile has left the propelling gun, the missile traveling along a settrajectory misses the target and is wasted.

In the use of the present invention, the chance of hitting a targetafter a change of course thereof, although fortuitous, is enhanced bythe firing of the steering charge which alters the course to apredetermined angle such, for example, as five degrees from theoriginal-trajectory. If the course of a missile in flight does notchange upon a change upon a change in course of a moving target, themissile will continue along its original set trajectory and wouldunquestionably miss the target, the kill possibility being nil. Inaccordance with the arrangement presented by the present invention, theprobability of the missile coming in proximity to the target isincreased by a ratio of l in 36 over that expected with a fixedtrajectory missile.

To achieve this enhanced probability of kill over that expected withfixed trajectory missiles, exhaust ports in communication with anexplosive steering charge are provided in the casing of the missilewhereby jet action imparts a transverse thrust thereto, while in flight,upon ignition of the charge. Since the invention employs a rotatingmissile, the transverse thrust may occur anywhere within the 360 rotatedby the missile, the angular spin position of the ports at the instant ofignition of a charge being fortuitous in determining the course of thenew trajectory taken by the missile. Since the angular orientation ofthe ports at the instant of ignition of the steering charge isfortuitous, there is a great probability that the ports may be sooriented as to direct the missile in a direction different from thatnecessary for target interception, resulting in a wasted missile as in aset trajectory missile; on the other hand, there is also theprobabilitythat the angular spin position of the ports may be such thatthe missile will be directed along a trajecory approaching the target.If the ports, at the instant of steering charge ignition, are positionedwithin a arc, the center ice of which arc is the exact angular spinposition for providing a collision course with the target, thetrajectories the missile may assume within this 10 are are such as tobring a missile in proximity to the target. Therefore, the probabilityof the ports being in a favorable angular spin position for the missileto assume a trajectory that approaches the target is 10 in 360, or 1 in36, or approximately 3 percent.

In the use of the steering charge, it is apparent that on a rightangular plane of the original trajectory of the missile at the targetinterception range there is a gradually decreasing spiral path of pointsto which the missile may be directed by firing the steering chargeduring the flight of the missile. If the steering charge is firedshortly after the missile leaves the muzzle of the gun, the correctivedistance on the aforementioned plane of target interception at rightangles to the original trajectory will be great, the distance becomingincreasingly less in the spiral as the firing of the steering charge isdelayed during the flight of the missile.

The missile of the present invention is rendered capable of operating inthe foregoing manner by the use of the system which will be hereinaftermore fully disclosed. The missile is provided with canted stabilizerfins set at an angle suflicient to impart a rotational speed ofapproximately ten revolutions per second during the flight thereof. Themissile is also provided with a modified proximity fuze which has meansfor transmitting an asymmetrical signal of predetermined frequencyrather than the usual symmetrical signal employed to fire the highexplosive charge as the missile comes into proximity with a target, thesignal being rendered asymmetrical for the purpose of indicating theinstantaneous rotational position of the missile in flight and beingadapted for reception at the launching point and also reflected from thetarget according to the conventional operation of the proximity fuze. Asis well known to those skilled in the art, the signal reflected from thetarget and received by the fuze is of a frequency different from thetransmitted predetermined frequency by an incremental amount due to theDoppler frequency shift effect. The signal transmitting means of thefuze of the present invention has the multiple functions of transmittingthe asymmetrical signal toward the target for reflection back to thefuze, transmitting the signal to the launching point of the missileWhere it may be received by suitable apparatus to indicate therotational position of the missile, receiving a signal from thelaunching point for firing the steering charge, and receiving thereflected signal thereby to fire the high explosive charge when thereflected signal reaches a predetermined intensity.

An object of the present invention is to provide a new and improvedsteering control method and system for a gun launched rotating missilehaving the feature of effectively altering the trajectory of the missileduring the flight thereof.

Another object of the present invention is to provide a new and improvedmethod and system of reaction steering control of a missile while inflight.

Another object is to provide a missile having a steering charge mountedtherein and capable of producing a known lateral thrust whereby thetrajectory thereof may be altered to a predetermined angle at a selectedtime during the flight of the missile.

Still another object of the invention is to provide a modified firingcircuit for a proximity fuze having provision for firing the reactionsteering charge when a target-simulated signal is transmitted to thefuze.

A further object is to provide a course changing missile capable ofbeing used with any conventional firing control system to increase theprobability of moving target interception above that expected with afixed trajectory missile.

Still another object is to provide a steering control arrangement for agun launched missile having suflicient ruggedness to withstand theforces attendant thereto.

A further object is to provide a control arrangement for altering thetrajectory of a missile in flight which is compact and of economicalconstruction.

Other objects and many of the attendant advantages of this inventionwill be readily appreciated as the same becomes better understood byreference to the accompanying drawings wherein:

a FIG. 1 is a view in diagrammatic form of the missile and steeringcontrol system of the present invention;

FIG. 2 is a schematic longitudinal sectional view of the missile of thepresent invention;

FIG. 3 is a sectional view taken along the line 3-3 of FIG. 2;

FIG. 4 is a fragmentary elevation of the tail portion of the missile andillustrating the canted stabilizing fins;

FIG. 5 is a diagram of the electrical system employed in the missile ofthe present invention;

FIG. 6 is a diagram of the radiation pattern as received at thelaunching point whereby the rotative position of the missile is madeknown;

FIG. 7 is a diagram indicating the altered trajectories which may beassumed by the missile during any predetermined one second of the flightof the missile; and

FIG. 8 is a diagram indicating the spiral path corresponding to thetrajectories illustrated in FIG. 7.

Referring now to the drawings wherein like reference characters indicatelike parts throughout the several views and more particularly referringto FIGS. 2, 3 and 4, the numeral 10 indicates generally a missileconstructed in accordance with the present invention. The missile 10comprises a casing 11 having'a modified proximity fuze 12 mounted in thenose portion thereof, the nose portion being insulated from casing 11and fuze 12 by suitable dielectric means. At the tail portion of themissile, stabilizing fins 13 are mounted, the fins being canted slightlyto rotate the missile at a speed of approximately 10 revolutions persecond during the flight thereof. There is mounted in the tail portion ahigh explosive charge 14. An electroresponsive detonator 15 is providedfor firing the charge 14.

Intermediate the nose and tail portions, a steering charge 16 isarranged within the casing 11 and balanced at the center of mass of themissile. The charge 16 comprises a plurality of groups of rocket typeexplosive discs 17, metallic supporting discs 18 being interposedbetween the groups of discs 17. In the wall of the casing 11 andadjacent each of the groups of explosive discs 17 there is provided aport 19. If desired, a plurality of the ports 19 may be provided foreach of the groups of explosive discs. It is, of course, understood thatall of the ports 19 are to be positioned on one side of the casing 11.

Since the missile rotates about 10 r.p.s., the resultant rotation of theangular position of the rocket ports 19 allows the lateral rocket thrustand hence the direction of trajectory change to be made in any directionin the plane normal to the trajectory.

Arranged adjacent the charge 16 at the nose end thereof is an igniter 21of the electroresponsive type. Bulkheads 22 and 23 separate thecompartment containing charge 16 from the tail compartment and the nosecompartment, respectively.

The proximity fuze 12 is arranged in the usual manner to ignite thedetonator 15 and thereby to fire the high explosive charge when themissile has reached a predetermined distance 'from the target.

In accordance with the invention and as will be more fully describedhereinafter, the transmitted radiation pattern of the fuze isasymmetrical in shape and, therefore, it is possible by the use of asuitable receiver at the launching pointto be informed of the exactangular or rotational position of the ports 19 at all times throughoutthe flight of the missile.

In the diagram of FIG. 5 there is illustrated a modified form ofproximity fuze circuit of the Doppler type as employed in the presentinvention and having a coil 32 mounted transversely in the nose portionof missile 10, the ends thereof being connected respectively to a pairof radiation ears 33 and 34 which are insulatably mounted on the noseportion of the missile to form a transverse dipole. A triode tube 35 hasthe grid 36 thereofconnected to the coil at one side of the centerthereof as at 37 while the coil is grounded to the casing 11 at theopposite side of the center thereof as at 38. Due to the connection ofcoil 32 to casing 11 at 38, the casing 11 functions as an end-fed dipoleand is axially excited by oscillations appearing across coil 32, thefrequency of oscillations being determined by the inductance of coil 32and the interelectrode capacitance of tube 35.

As a result of the arrangement of a pair of dipoles disposed as toradiate fields at right angles to each other and as a result of theantennae connections of coil 32 whereby the dipoles are simultaneouslyexcited, the combined effects of these perpendicularly radiated fields,when viewed by a vertically polarized antenna at the launching point,results in the reception at the launching point of the aforementionedasymmertrical radiation pattern. This asymmetrical radiation pattern maybe received at the launching point on .a vertical dipole antenna to givea signal modulated in the form illustrated by the envelope wave 39 ofFIG. 6. The asymmetrical envelope wave 39 has high points 41 and lowerpoints 42 which occur when the cars 33 and 34 respectively are facingtoward the earth. The steering charge ports 19 are oriented withrelation to the ears 33 and 34 in such a manner that, for example, theports are positioned on the same side of the missile as the ear 33thereby providing a means of informing the operator at the point oflaunching of the instant rotational position of the missile at all timesthroughout the flight thereof. As is apparent from this arrangement andfrom envelope 39 of FIG. 6, the signal strength which is received at thelaunching point is proportional to the orientation of the missile or,what amounts to the same thing, is proportional to the angularorientation of the steering ports. When the ports 19 are up, maximumsignal strength, as indicated at 41, is received by the receiver; and,when the ports are facing the earth, a signal strength indicated at 42,which is considerably less than the maximum signal strength, is receivedat the launching position.

The tube 35 is provided with a filament type cathode 43 the circuitthereof comprising a choke 44 to ground on one side and a choke 45 andbattery 46 to ground on the other side thereof.

Plate 47 of tube 35 is connected by coupling condenser 54 to an audioamplifier 48, an.A.C. by-pass condenser 51 being interposed betweenplate 47 and ground and a plate load resistor 52 and B battery 53 alsobeing interposed between plate 47 and ground.

. The output side of amplifier 48 is connected by a C battery 58 to thegrid 59 of thyratron 60. Plate 61 of thyratron 60 is connected by aresistor 62 to battery 63 which is connected to ground on the other sidethereof. A condenser 64 is connected from the grounded side of battery63 to the plate 61. The cathode 65 of thyratron 60 is connected toground through detonator 21 for the steering charge in parallel with anexplosive switch 66 which may be of any type suitable for the purposesuch, for example, as the explosive switch disclosed in the copendingapplication of Howard C. Filbert, Jr. for Explosive Operated'PressureSwitch, Ser. No. 130,821, filed Dec. 2, 1949, now US. Patent No.2,721,240 which issued on Oct. 18, 1955. Cathode 65 is likewiseconnected in open circuit arrangement with detonator 15 for the highexplosive charge through the initially open contacts 55 of which 66. Asthe steering charge is fired, the contacts 55 of explosive switch 66 areclosed, thereby connecting the cathode 65 to the detonator 15. It willthus be seen that the first pulsation of tube 60, which pulsation is ofshort duration, fires the igniter 21 thereby firing the steering chargeand simultaneously causing the explosive switch 66 to close the circuitto the detonator 15. The second pulsation of tube 60 fires detonator 15,thereby firing the explosive charge 14.

In the event that the target does not change course and consequently thetrajectory of the missile has not been changed, the steering charge willbe detonated as soon as the missile is sufiiciently close to the targetto render thyratron 60 conductive, and almost instantaneously thereafterthyratron 60 will again be rendered conductive to detonate the explosivecharge 14, since the charging time of condenser 64 is very short. Inview of the fact that the two conductions of thyratron 60 aresuccedently instantaneous and in view of the fact that the missile isvery close to the target at the time thyratron 60 initially' becomesconductive, it is evident that the missiles course is not suflicientlydeflected as to miss the target.

In the operation of the circuit, the output of tube 35, asaforedescribed, produces the asymmetrical wave pattern 39 by reason ofthe simultaneous radiation from the casing excited axially as a dipoleand from the transverse dipole formed by cars 33 and 34.

The receiver 67 at the launching position may pick up the asymmetricalradiation at a predetermined frequency F which is the frequencydetermined by the inductance of coil 32 and the interelectrodecapacitance of tube 35, and may be adapted to feed it to a fire controlunit 68, not a part of this invention, which unit, at the prescribedtime, as determined by the information retained in the fire controlunit, may actuate transmitter 72 at any suitable instant to transmit atarget-simulating signal of short duration, the frequency of thetarget-simulating signal being F -I-AF where AF is a preassignedincremental frequency shift simulating the Doppler effect frequencyshift introduced in the signal reflected by the moving target. If nofire control system is used, any suitable manually operated transmittermay be employed to transmit the desired signal. This transmittedsignalmodulates the oscillator 35 to derive the difference frequency AF whichis applied to audio amplifier 48, amplifier 48 having a narrow pass bandcentered at AF. Since the limits of velocity range of aircrafts areknown and since the velocity of the missile is known, the range ofrelative velocities between the targets and missile can be calculated,and from these calculations the various Doppler frequency shiftsaffecting a radiated signal of predetermined frequency and occurringwithin the range of relative velocities can be determined. From thisdetermination, the pass band of amplifier 48 is designed so as to passsubstantially all incremental frequencies caused by the Doppler effectwith the aforementioned range of relative velocities, the centerfrequency of this pass band being AF.

Amplifier 48 amplifies the passed signal AF and ap plies it to the grid59 of thyratron 60 which is normally maintained at a non-conducting biasby battery 58. The signal transmitted by transmitter 72 is designed tobe of sufficient strength to overcome the biasing potential of battery58 and trigger the thyratron 60.

When the thyratron is rendered conducting, the energy stored incondenser 64 is discharged therethrough by way of plate 61, cathode 65,and thence through the explosive switch 66 in parallel with igniter 21to ground, thereby to fire the steering charge 16.

Upon operation of switch 66, cathode 65 is in closed circuit arrangementwith detonator 15 for the high explosive charge 14. As voltage stored incondenser 64 is only suflicient to fire the explosive switch and igniter21 and is of short duration, the detonator 15 is not fired during thefirst signal. As condenser 64 is discharged, the voltage on plate 61 ofthe thyratron is reduced below the voltage for sustaining conduction andthe tube is extinguished. Condenser 64 is charged from battery 63through resistor 62 which has a resistance sufiicient to reduce theplate potential as aforementioned while also providing for rapidcharging of the condenser.

As the missile approaches the target, the radiation pattern of frequencyF transmitted by the missile is reflected and shifted in frequency anincremental amount AF by the target, the signal reflected back to themissile assuming the frequency F +AF where AF corresponds to the Dopplerincremental frequency shift caused by the velocity of the aircraft. Thereflected signal F +AF is mod-ulated in oscillator 35 to derive thedifference frequency AF which is passed by amplifier 48 and applied togrid 59 of thyratron 60. The bias 58 is selected to have a value such asto maintain thyratron 60 insensitive to the reflected signals, whichincrease in intensity as the missile approaches the target, until themissile is within a predetermined distance, for example 40 feet, fromthe target at which distance the reflected signal is of such intensityas to produce a signal in the output of an amplifier having an amplitudesufiicient to overcome the bias on thyratron 60.

When the amplitude of the output of amplifier 48 in response to thereflected signals reaches a predetermined value to overcome the bias 58as described in the preceding paragraph, the firing circuit includingthyratron 60 operates in the manner heretofore described with theexception that the voltage of condenser 64 is applied to detonator 15,the switch 66 being closed. It will thus be seen that the circuit isarranged in such a manner that on receiving a first signal from thelaunching point the course of the missile is changed and on receiving asecond signal (the reflected signal) the main explosive charge is fired.

Referring particularly to FIG. 1 in which the system for operation ofthe missile of the present invention is illustrated, the missile 10 isfired from a smooth bore gun 24 along a trajectory A in order tointercept target 25 moving on a course B. Since the projectile travelswith a finite speed, the time of projectile flight must be allowed forand the gunis aimed at a point in space ahead of the aircraft on apredicted flight path computed from the targets measured range andvector velocity detected by director 69, resulting in a predictedcollision point where course B and trajectory A intersect. At this pointof intersection, there is a gradually decreasing spiral path of pointsin a plane normal to trajectory A to which the missile may be steered,if desired, by firing the steering charge.

Assume that, after firing the missile 10, target 25 changes its course,for example, to course C. In order to change the course of the missileto intercept target 25 at point E, the steering charge 16 of the missilemust be fired when the missile reaches the proper rotational positionand the proper position along the trajectory to cause such interceptionby a five degree change of course. Firing of the steering charge iseffectuated at random by transmitting from transmitter 72 a signal ofshort duration which, in turn is received by the missile and fires thesteering charge, as aforedescribed. If reception of this signal by themissile occurs when ports 19 are within an arc of 5 degrees on each sideof the proper rotational position for attaining course D, then themissile will assume a trajectory approximating course D, as hereinabovedescribed, and eventually comes Within proximity of the target nearintercept point E.

Referring to FIGS. 7 and 8, there is illustrated the selectivetrajectories or altered courses which may occur during one second of theflight of the missile. As will be apparent in FIG. 7, the missile willmake 10 revolutions during this period of'time, and the steering chargewill be aligned to change the course of the missile in the 0 or 360position 10 times. The selective trajectories move in a spiral towardthe original trajectory, the distance between the convolutions being, intime, one tenth of a second, the forward motion of the missile causingthe spiral to become of increasingly smaller diameter as the missileapproaches the target interception range. It will, thus be seen that ifthe steering charge is fired shortly after leaving the gun and at agreat distance from the target interception range the spiral will belarge and the deviation from the original trajectory great, while if thesteering charge is fired at a point near the target interception range,the spiral will be small and the deviation from the original trajectoryless. It will, also, be understood that while an angle of five degreesof deviation has been described herein, any desired angle of deviationmay be employed by changing the size or force of the steering charge.

It is to be understood that in employing a steering charge of a givenforce there will be a slight variation in the exact angle of deviationas the forward speed of the missile is reduced during the flightthereof.

It is also to be understood that the present invention is directed to acourse-changing missile, per se.

Obviously many modifications and variations of the present invention arepossible in the light of the above teachings. It is therefore to beunderstood that within the scope of the appended claims the inventionmay be practiced otherwise than as specifically described. I

What is claimed and desired to be secured by Letters Patent of theUnited States is:

1. Apparatus for changing the course of a missile in flight comprising,a casing for said missile, a reaction charge disposed within said casingand arranged to direct an impulse through one side of said casing andthrough the center of gravity of the missile as the charge is firedwhereby the missile is given a lateral thrust and is directed along anew trajectory with the longitudinal axis of the missile initially at anangle with respect thereto, means on the missile for bringing said axisthereof into alignment with the new trajectory, and means carried by themissile and responsive to a transmitted signal from the launching pointfor firing said charge.

2. A missile comprising, a casing having a port in one side thereof, areaction charge mounted within said casing and adjacent said port andarranged to direct an impulse therethrough as the charge is firedwhereby the course of the missile is changed, and means including atransmitter located at the launching point of the missile for firingsaid charge during the flight of the missile.

3. Apparatus for changing the course of a missile in flight comprising,a casing for said missile, canted fins arranged on the tail portion ofthe missile for maintaining alignment of the longitudinal axis thereofwith the trajectory and for causing rotation of the missile about saidaxis at a predetermined rate, a reaction charge disposed within saidcasing and arranged to direct an impulse through one side of the casingand through the center of gravity of the missile as the charge is firedwhereby the missile is first moved along a new trajectory at an anglethereto and thereafter is brought into alignment there- 8 with by saidfins, means in the missile for firing the charge, and means including atransmitter at the launch ing point for igniting said charge firingmeans while the missile is in flight.

4. Apparatus for performing a course-changing operation on a guidedmissile subsequent to the launching thereof comprising, a substantiallytubular casing for said missile having a plurality of perforations inone side thereof, a reaction steering charge mounted within said casingand arranged to direct a reaction impulse through the perforationsthereof as the charge is fired, and means including a transmitter at thelaunching point for firing said charge during the flight of the missile.

5. In a control system for a gun-launched missile, the combination of anexplosive charge disposed within the missile, a reaction steering chargedisposed at the center of mass of the missile and arranged to direct areaction force laterally thereof as the steering charge is fired, cantedstabilizing fins on the missile for imparting rotation thereto, aproximity fuze constructed and arranged to transmit an asymmetricalradiation pattern of energy suitable for reflection from the target, andelectroexplosive switch means operable concurrently with the firing ofthe steering charge for connecting the explosive charge in paralleltherewith as the steering charge is fired, said fuze having means forfiring said steering charge and said electroexplosive switch means inresponse to energy of predetermined frequency received thereby andtransmitted thereto from the launching point, said last named meansbeing arranged to fire said explosive charge in response to saidreflected energy as the missile moves into predetermined spaced relationwith respect to a target after the steering charge has been fired.

References Cited BENJAMIN A. BORCHELT, Primary Examiner.

JAMES L. BREWRINK, NORMAN H. EVANS,

' Examiners.

A. GAUss, J". w. GALLAGHER, H. G. WEISSEN- BERGER, v. R. PENDEGRASS,

Assistant Examiners.

1. APPARATUS FOR CHANGING THE COURSE OF A MISSILE IN FLIGHT COMPRISING,A CASING FOR SAID MISSILE, A REACTION CHARGE DISPOSED WITHIN SAID CASINGAND ARRAGED TO DIRECT AN IMPULSE THROUGH ONE SIDE OF SAID CASING ANDTHROUGH THE CENTER OF GRAVITY OF THE MISSILE AS THE CHARGE IS FIREDWHEREBY THE MISSILE IS GIVEN A LATERAL THRUST AND IS DIRECTED ALONG ANEW TRAJECTORY WITH THE LONGITUDINAL AXIS OF THE MISSILE INITIALLY AT ANANGLE WITH RESPECT THERETO, MEANS ON THE MISSILE FOR BRINGING SAID AXISTHEREOF INTO ALIGNMENT WITH THE NEW TRAJECTORY, AND MEANS CARRIED BY THEMISSILE AND RESPONSIVE TO A TRANSMITTED SIGNAL FROM THE LAUNCHING POINTFOR FIRING SAID CHARGE.